Ever more components of fiber composite materials are used in the construction of vehicles and aircraft. These fiber composite materials preferably consist of glass, carbon or aramid fibers, which are built-up of fiber layers and connected with one another through polymer materials. In that regard, the components are generally produced through lamination of the synthetic resin impregnated fiber layers under pressure and heat in a pressing form or mold. These composite materials are usually lighter than comparable metal structural parts and possess a high stiffness and strength and are therefore preferably utilized in aircraft construction.
For such structures, the exact documentation of the operational loads is of great interest, in order to prove or document the remaining operating life of the structure as near as possible to actual reality. In this manner, the permissible operating times for each structure can be fully utilized in an economically optimal manner. For that purpose it is primarily necessary to record and to document the prevalence and the magnitude of the structural deformations. These are summarized to load collectives and correlated with the determined operating life of the material. Thereby, on the one hand the maintenance and service intervals and on the other hand the remaining operating life can be adapted to the conditions that actually arose during operation, and thus the structure can be operated in an economically optimal manner.
Moreover, damages such as cracks or delaminations can arise in the material of lightweight structures due to high loads or through too-high production tolerances, such as shrinkage voids or sink holes or fiber proportions or components. These damages can considerably weaken the mechanical stiffness and the strength of the components. Especially in connection with aircraft, such components are also subjected to the danger of impact damages through birds and ice particles during operation. These loads can lead to previously mentioned damages within the composite materials, which are not externally visible and represent an endangerment of safety. In order to be able to detect or determine such damages, it is known to recognize these in the regularly occurring maintenance procedures, through non-destructive testing methods such as x-ray or ultrasound tests. However, in that regard the danger exists, that a clear reduction of the operating strength arises until the time of the next maintenance inspection due to damage growth as a result of high vibration or oscillation loads, whereby such reduction of the operating strength is to be avoided in all cases. Therefore, a series of possibilities exists, to detect such dangers immediately especially on aircraft components, in order to remove or correct the damages as early as possible.
Often, however, it is also necessary to inspect such components or other fiber composite material components before the installation or utilization in the aircraft construction and the like, in order to determine the operating life and to avoid constructive damages tending to cause danger. In that regard it is necessary to impose prescribed loads on the components that are to be inspected or examined, in order to prepare a proof or documentation of the operating life and to determine strains tending to cause damages on the components and to recognize a danger of damage at an early time. In that regard, the method for the monitoring and for the examination primarily distinguishes itself through the evaluation and in the attainment of the measuring results in the form of a loading analysis.
An apparatus for the determination of impact damages on fiber composite material components is known from the DE 40 25 564 C1. For that purpose, a plurality of distributed arranged piezoelectric foil elements are secured on an outer surface side of the vehicle body components that are usually only a few millimeters thick, and are lead to an electronic monitoring arrangement via an electrical connection. Upon the occurrence of a strong compression influence through an impact loading, which can lead to a delamination, a capacitive charge variation arises in the piezo transducer elements arranged in the proximity, whereby the capacitive charge variation is essentially proportional to the impact pressure. This charge variation is then detected in a monitoring arrangement and can be indicated corresponding to the damage-relevant impact pressure and location, in order to immediately introduce a targeted damage examination. However, with such a monitoring apparatus, only excessive impact loadings that can lead to a delamination are detectable. An exact documentation of the operating loads for the evidence or proof of the remaining operating life, as well as damages on the fiber composite material components that arise through other excessive strain loads that are not dependent on pressure, are not detectable with this monitoring apparatus. Especially, with such a monitoring apparatus for examination purposes, only impact loadings on prescribed construction parts can be analyzed.
It is known to apply an optical reflection diffraction grating on a fiber composite material, from the DE 35 20 664 A1. For monitoring the surface strain, the reflection grating can be illuminated with a laser light beam, and the radiation intensity thereof in a certain reflection direction can be detected. If the surface of the material changes due to a strain expansion or compression, thereby the diffraction angles and thus also the radiation intensity in the detected directions also change. Such a radiation intensity is then measured with opto-electronic position detectors, and can be indicated as a value of the surface strain. Such a monitoring of the material surfaces is, however, only possible where this surface can be radiated with laser light and the reflected or re-radiation intensity thereof is detectable at a certain spacing distance from the surface. Especially when the surfaces are additionally provided with other protective or insulation layers, which do not follow the strain, then such a monitoring or a strain examination cannot be carried out.
The detection of a surface strain with strain gages on a rotational shaft of a fiber composite material is known from the DE 40 21 914 C2. This strain measurement is carried out for the rotational moment or torque determination by means of a testing machine, whereby the strain gages are glued or adhesively applied onto the surface of a fiber composite pipe, and the strain detection thereof serve for the calculation of the torque in the torsional body. For that purpose, apparently commercially available typical strain gages are used, which are not suitable for measurements of strain tending toward damage on fiber composite material surfaces, because typical commercially available measuring grids do not withstand such strain regions. Moreover it is not known, that the determination of the rotational moment or torque is utilized for the proof or evidence of the remaining operating life of the component.
Therefore, such typical commercially available strain gages would have to be renewed after each examination test run or each loading with surface strains tending to cause damage, which brings about a considerable cost-intensive expense and effort especially in connection with multi-point measurements. Especially, in material analysis measurements, no measured values could further be evaluated in the upper damage-inducing range, so that thereby also only an inadequate analysis result can be achieved. While it is conceivable to produce special strain gages of wire measuring grids for such surface tension measurements, which wire measuring grids also withstand greater strain ranges on composite fiber surfaces, which would be uneconomical, however, for multi-point measurements for component analysis or for the monitoring of large surface area aircraft parts.
Strain gages and methods for their production are, however, previously known from the EP 0 667 514 B1. These basically consist of a typical commercially available photolithographically produced measuring grid that is vapor deposited on a carrier film and additionally is covered with a protective layer. For the connection, this measuring grid has flat soldering connection surfaces, which represent the beginning and the end of the measuring grid. Connection wires are soldered thereon for the wiring, and are lead to the provided connection parts for the circuit-connection. Such a strain gage can basically be applied only on the surface of a strain body, because otherwise a subsequent wiring-connection is no longer possible. A previous wiring-connection would also be unrealistic, because an economical handling of a plurality of connection wires is hardly possible in the known production methods of composite materials.